The present invention relates to small propulsion systems for maneuvering space craft and, more particularly, to electrothermal arcjet thrusters having an improved anode.
An electrothermal arcjet thruster converts electrical energy to thermal energy by heat transfer from an arc discharge to a flowing propellant and from thermal energy to directed kinetic energy by expansion of the heated propellant through a nozzle. The electrothermal arcjet thruster is described in more detail in U.S. Pat. No. 4,548,033, to Cann, U.S. Pat. No. 4,800,716 to W. W. Smith et al., and U.S. Pat. No. 4,926,632 to R. D. Smith et al., all of which are incorporated, by reference, in their entireties herein.
Most electrothermal arcjet thrusters have as common features an anode in the form of a nozzle and a cathode in the form of cylindrical rod with a conical tip. The anode has an arc chamber defined by a constrictor in a rearward portion of the body and a nozzle in a forward portion thereof. The cathode rod is aligned on the longitudinal axis of the anode with its conical tip extending into the upstream end of the arc chamber in spaced relation to the constrictor so as to define a gap therebetween.
An electric arc is initiated between the cathode rod and the anode at the entrance to the constrictor. The arc is forced downstream through the constrictor by pressurized flow of a propellant gas. The arc stabilizes and attaches at the nozzle. The propellant gas is heated in the region of the constrictor and the region of arc attachment at the mouth of the nozzle. The superheated gas is then exhausted out the nozzle to achieve thrust.
Historically, pure propellants, typically ammonia (NH.sub.3) or hydrogen (H.sub.2) have been used in electrothermal arcjet thrusters. More recently, hydrazine (N.sub.2 H.sub.4) has been used as a propellant in arcjet thrusters. Propellants such as ammonia and hydrazine are storable in space as a liquid without refrigeration, while cryogenic propellants such as hydrogen and helium are not.
The 1500-2000 watt arcjet base propulsion devices are typically limited to approximately 550 seconds of specific impulse. Specific impulse is defined as the generated thrust (in pounds force) divided by the propellant consumption (pounds mass per second). Beyond a specific impulse of 550 seconds, the constrictor diameter reduces deteriorating thrust performance.
Applicant has determined that the reduction in the diameter of the constrictor is due both to the erosive action of the arc upstream of the constrictor during initiation and to the thermal gradient between the constrictor region and nozzle body radially outward from the constrictor during operation. While means have been proposed to limit the erosive effect of the plasma gas, as in U.S. Pat. No. 2,862,009 to Gage, which discloses a shielding gas passing between the plasma gas and the nozzle to keep the plasma flame from impacting the nozzle, such solutions have not proved adequate to sufficiently enhance the performance of spaceship propulsion units.